The advantages of composite structures in aircraft construction are well known in the aerospace industry. Composites are increasingly used in the primary structure of new aircraft, both as an alternative to and adjunct to metal structures. Carbon fiber composites in which carbon fibers are embedded in an epoxy matrix offer relatively high stiffness and high strength while maintaining a low weight.
In one manner of producing large commercial transport aircraft the fuselage structure consists of a thin load bearing skin supported by lateral stringers to eliminate buckling and circumferential bulkheads to transfer shear and retain the fuselage shape. In metal aircraft construction, bulkheads are commonly joined to the skin by the use of rivets. In some large composite aircraft, including the prior art Boeing™ 787, these bulkheads are joined to the skin with an intermediary clip, necessitating the use of two rows of fasteners, the first between the clip and skin, and the second between the clip and bulkhead. While this method is effective at joining the bulkhead to the skin, it results in an excessive quantity of fasteners, and carries with it significant weight, complexity, and cost penalties.
Fuselage and wing skins in aircraft are commonly load bearing in both metal and composite construction. The skins provide a large surface area through which to dissipate shear and bending loads. This large thin structure is susceptible to buckling; therefore an internal support structure is required to adequately distribute loads to the skin and to prevent these skin panels from buckling. Major load carrying components such as wing mounts or landing gear are attached to internal frames which then distribute their concentrated loads out to the distributed skin panels. There are several prior art arrangements of the composite skin and supporting internal structure.
Relatively smaller composite aircraft such as the Raytheon™ Premier have used a load carrying skin structure which is supported against buckling by a lightweight honeycomb core. The use of a lightweight core material between skins is referred to as sandwich construction. FIG. 1 shows the nose fuselage section 100 of a Raytheon™ Premier. An outer skin layer 110 and an inner skin layer 120 surround a honeycomb core layer 130.
Larger composite aircraft such as the Boeing™ 787 have used the aforementioned skin and stringer configuration. In this configuration, stringers support the skin and major loads are transferred to the skin through the bulkheads spaced longitudinally through the aircraft. FIG. 2 depicts a Boeing™ 787 fuselage barrel section 200, comprising a skin 210 supported by regularly-spaced longitudinal stringers 220 and circumferential bulkheads 230. Fasteners 240 are used in attaching the bulkhead to the skin.
Other structural construction variations are also possible. In a hybrid configuration, bulkheads are used in combination with cored skin sandwich structure which is interrupted or tapered down at each bulkhead interface. In this configuration, the core prevents buckling of the skin and the bulkhead transfers concentrated loads to the skin.
Aircraft composite materials must often be cured to obtain the desired properties. Curing usually involves exposing the structure to combinations of one or all of elevated temperature, elevated pressure, or diminished pressure. A composite structure is considered “co-cured” when all the layers or components of the structure are cured together in a single curing stage, even if some of the layers or components were exposed to some type of curing before the step of co-curing. Co-curing can result in very strong bonds between parts and composite layers. In recognition of this, governmental civil aircraft certification agencies including the FAA currently approve of such co-cured structure without additional riveting between the skin and stringers.
Composite structures are often built of assemblies of co-cured parts. Such assembly uses secondary bonding. In compliance with current government certification practice, this secondary bonding between parts takes the form of rivets.
Recent composite aircraft fuselage or wing construction as found on the prior art Boeing™ 787 uses co-cured outer skin 210 and stringers 220 in the construction of a fuselage section 200, which avoids the need for a high number of rivets to attach the stringer 220 to the skin 210. The stringers 220 are co-cured with the skin 210 and continuous on either side of the bulkhead 230. As a consequence, the circumferential bulkheads 230 must be contoured around the stringers 210 by means of cutouts 232. These bulkheads 230 (or ribs in the case of wing construction) are numerous in a typical transport aircraft. The assembly of bulkheads 230 from multiple composite parts on the Boeing™ 787 requires secondary bonding and a large number of rivets 240 for each bulkhead 230. Each bulkhead 230 is riveted to a series of L-shaped clips around its circumference; each clip in turn is riveted to the skin. FIG. 3 shows another view of a prior art Boeing™ 787 fuselage section 300. Floor beams 310, which separate between the passenger section 320 and the cargo hold section 330, are built separately and attached to each bulkhead 340 using rivets 350 as secondary fasteners.